The present invention relates generally to spacecraft control systems, and, more particularly, to spacecraft control systems that provide for control of thermal shock disturbance.
Many spacecraft, such as earth orbiting communication satellites, require a particular attitude relative to the earth. Maneuvers to maintain earth pointing should take minimal time because loss of earth pointing often means that the spacecraft is not performing mission objectives. Many maneuvers to maintain earth pointing currently require torques supplied by thrusters. However, thrusters use limited fuel and thereby potentially shorten the life of the spacecraft mission.
Currently, in conjunction with thrusters, many spacecrafts use flywheels and reaction wheels as momentum wheels to control attitude. Momentum wheels are also used as energy storage mechanisms, which provide power to the spacecraft, and thereby minimize use of heavy chemical batteries.
Thermal shock disturbance is an industry-wide problem experienced by earth orbiting spacecraft with solar wings. When a spacecraft enters an eclipse, abrupt temperature changes deform or bend the solar wings. This bending exchanges momentum between the body of the spacecraft and the wings which results in short-term attitude error. The severity of the attitude error depends primarily on the size of the solar wing. More specifically, thermal shock disturbance occurs in two situations. First, when the sun shines on a solar wing, the wing tends to bend away from the sun. Second, when the sun is not shining on the wing, the wing tends to bend toward the sun. As the solar wing moves in a particular direction, the spacecraft body moves in the opposite direction, according to the momentum conservation principle. In other words, thermal shock is a thermal-elastic effect where eclipse of the solar wing from sunlight causes the temperature to decrease rapidly because energy is subsequently radiated into deep space. The temperature change causes thermal deformation (deflection) of the wing. The deformation causes motion of the wing and thus motion (momentum change) of the central body.
Conventional solutions for thermal shock disturbance fall into two general categories. The first category relates to the mechanical design and placement of the solar wing. For example, many prior art solutions have concentrated on modeling the thermal shock phenomenon and designing solar panels to minimize momentum exchange. Also, innovative mounting of the solar wing may reduce momentum exchange. However, these solutions tend to significantly increase the cost of the spacecraft. The second solution relates to the design of internal spacecraft control systems.
Because of unwanted motion of the body, the spacecraft is pointing away from its target, and, subsequently, control systems are often used to correct and maintain the desired pointing. Feed-forward systems are often constructed to deal with thermal shock. However, feed-forward systems are limited by the ability of their designers to construct a signal that sufficiently compensates for the given thermal shock. Position-dominated feedback controllers, such as PID (proportion, integral, derivative) controllers, are also typically used in satellite control systems. PID controllers are generally more robust for maintaining the necessary earth attitude than feed-forward systems. However, typical PID controllers tend to have long transient times before reaching steady-state.
The disadvantages associated with these conventional spacecraft design and control system techniques have made it apparent that a new technique for minimizing thermal shock is needed. The new technique should have rate-dominated high-bandwidth feedback controller and should include a spacecraft actuator design that efficiently responds to the feedback controller. The present invention is directed to these ends.
It is an object of the present invention to provide an improved thermal shock suppression system. It is also an object of the present invention to provide an improved thermal shock suppression system for a satellite.
In accordance with the present invention, a satellite system, which includes a solar wing moveably connected to a satellite central body, is disclosed. A sensor, also coupled to the satellite central body, detects the movement of the body and generates a rate signal based on that movement. Additionally, an actuator, which controls momentum, is coupled to the satellite central body with maximum torque along the thermal shock axis. Subsequently, a rate-dominated thermal shock suppression controller, which is coupled to the satellite central body, receives the rate signal from the sensor to control the actuator.
Additional advantages and features of the present invention will become apparent from the description that follows and may be realized by the instrumentalities and combinations particularly pointed out in the appended claims, taken in conjunction with the accompanying drawings.